The present invention relates to a turbine engine, and more particularly to an improved compressor for a tip turbine engine.
An aircraft gas turbine engine of the conventional turbofan type generally includes a forward bypass fan and a low pressure compressor, a middle core engine, and an aft low pressure turbine, all located along a common longitudinal axis. A high pressure compressor and a high pressure turbine of the core engine are interconnected by a high spool shaft. The high pressure compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in a combustor, where it is ignited to form a high energy gas stream. The gas stream flows axially aft to rotatably drive the high pressure turbine, which rotatably drives the high pressure compressor via the high spool shaft. The gas stream leaving the high pressure turbine is expanded through the low pressure turbine, which rotatably drives the bypass fan and low pressure compressor via a low spool shaft.
Although highly efficient, conventional turbofan engines operate in an axial flow relationship. The axial flow relationship results in a relatively complicated elongated engine structure of considerable length relative to the engine diameter. This elongated shape may complicate or prevent packaging of the engine into particular applications.
A recent development in gas turbine engines is the tip turbine engine. Tip turbine engines locate an axial compressor forward of a bypass fan which includes hollow fan blades that receive airflow from the axial compressor therethrough such that the hollow fan blades operate as a centrifugal compressor. Compressed core airflow from the hollow fan blades is mixed with fuel in an annular combustor, where it is ignited to form a high energy gas stream which drives the turbine that is integrated onto the tips of the hollow bypass fan blades for rotation therewith as generally disclosed in U.S. Patent Application Publication Nos.: 20030192303; 20030192304; and 20040025490. The tip turbine engine provides a thrust-to-weight ratio equivalent to or greater than conventional turbofan engines of the same class, but within a package of significantly shorter length.
Although much shorter axially than conventional turbine engines, much of the length of the tip turbine engine results from the number of stages in the axial compressor. Reducing the number of compressor stages would further decrease the axial length of the tip turbine engine.
The number of stages could be reduced by using larger chord compressor blades that do more work in turning and compressing the air. However, at some point, the compressor blade tends to separate from the blade and the blade becomes highly inefficient, and can result in engine stall.
Aspirated compressors have been used in conventional turbine engines to reduce the number of stages required in the compressor. In an aspirated compressor, suction is provided at selected locations on the surface of the compressor blades. The suction keeps the flow attached to the blade even with increased curvature and longer blade chord lengths. Aspirated compressors have not been implemented in tip turbine engines, which already have a shorter axial dimension.